Propellant composition and methods of preparation and use thereof

ABSTRACT

The invention relates to propellant compositions comprising a solid inorganic perchlorate oxidizing agent, a nitrogen-containing fuel, and a burn rate catalyst. Such compositions may be used as a propellant material, (e.g., in rocketry), a pyrotechnic material, an explosive material, a light generating material, a heat generating material, or a sound generating material.

CROSS-REFERENCE TO RELATED APPLICATION

This application is a continuation-in-part application of and claims the benefit under 35 U.S.C. § 120 to pending U.S. patent application Ser. No. 10/295,308, entitled COMPOSITE PROPELLANT COMPOSITIONS, filed Nov. 14, 2002, which is incorporated herein by reference in its entirety.

BACKGROUND OF THE INVENTION

1. Field of the Invention

This invention relates to propellant compositions as well as to methods of preparation and use of such compositions and, in particular, to composite propellant compositions suitable for solid rocket motor and gas generating applications.

2. Discussion of Related Art

Composite propellant compositions typically contain separate fuel and oxidizer components that are intimately mixed. Black powder, sometimes referred to as gunpowder is the oldest composite propellant composition. Black powder is made with charcoal, sulfur, and potassium nitrate (saltpeter). In black powder, potassium nitrate functions as the oxidizer, while sulfur and charcoal (i.e., carbon) are the fuel components. Even though black powder has a relatively low specific impulse, it has been used in rocketry and pyrotechnics for centuries.

Black powder rocket engines (rocket motors) are usually manufactured by pressing black powder into multi-layer paper casings under high pressure. The rear section of the engine is fitted with a nozzle through which exhaust gases escape and which is typically made from heat-resistant materials. In a typical rocketry application, an intimate mixture of 75% potassium nitrate, 15% charcoal, and 10% sulfur (by weight) is tightly packed into a case or casing, usually a multi-layered paper tube. An electrical igniter is used to ignite the rocket engine. Because the black powder is tightly packed to a uniform density, it can burn evenly and produces thrust as the hot expanding gases escape the rocket engine through the nozzle. Because black powder is granular and pressable, rocket engine production can be easily automated by means of multiple feed, hydraulic pressing machinery.

Although black powder is relatively inexpensive and readily available, it has a relatively low specific impulse. Other solid composite propellants have been developed using ammonium perchlorate, a fuel, and a binder. These composite propellants, sometimes referred to as AP composites or castable composites, can be formulated to produce greater energy and superior physical properties; however, these improvements are achieved at the expense of processing simplicity. By their nature, castable propellants are high viscosity liquids that can be cast (poured) into large diameter motor cases with relative ease; however, casting becomes more difficult as case dimensions decrease. Castable compositions also begin to cure the moment they are mixed; therefore, viscosity is time dependent and continues to increase until the composition can no longer be processed. In consequence, the casting process generates a substantial quantity of “waste” material.

The majority of castable propellant compositions are based on plural component, curable polymer systems consisting of a polyol (resin) and a diisocyanate (curative). Diisocyanates exhibit high rates of reaction with water; therefore, the first step in the manufacturing process is to reduce the water content of all propellant ingredients to near anhydrous levels, which in practice is less than 0.02%. Once the raw materials are dried, with the exception of the curative, all liquid ingredients are charged into a vertical vacuum mixer and mixed under vacuum for a prescribed period of time to remove dissolved gases and obtain a uniform blend of ingredients. This is expensive and time-consuming.

To achieve useful levels of energy and density, castable propellants are formulated in such a manner as to obtain high solids loadings (82% to 88%), which is defined as the weight percent solid ingredients to liquid ingredients. At these solids loadings, great care must be taken to preserve the fluidity, therefore, the castability of the mixture. This is accomplished by grinding solid ingredients (usually the oxidizer) to a series of predetermined and carefully controlled particle size distributions, thereby decreasing the void volume between solid particles. In doing so, the volume ratio of liquid to solids is maximized, which reduces the end-of-mix viscosity to manageable levels.

Once the solid ingredients are properly ground and sized, they are charged into the mixer in increments to avoid a potentially dangerous condition referred to as “dry mix.” Following the final ingredient addition, the mixture is mixed under vacuum for a specified period of time, usually between four and ten hours. Upon completion of the mix cycle, the curative is added and the entire composition mixed under vacuum for an additional hour. At this point, the propellant can be cast directly into motor cases or liners.

Due to the relative ease by which viscous liquids can entrap air bubbles, castable propellants must be introduced into individual motor cases under vacuum, or by bottom feeding by means of a casting bayonet. This is a complicated, capital equipment intense procedure that does not lend itself to high rate automation. In addition, the long cure time at ambient temperature (two to four weeks) results in an excessive amount of material and floor space allocated to “work in progress.” Beyond the financial disadvantage, the material being held as “work in progress” normally cannot be subjected to quality assurance testing until the final cure takes place.

Rocket engines using castable composites are typically assembled by machining the grain(s) to achieve the desired grain geometry, including the core, fitting the composite grain(s) into a special casing; inserting a nozzle through which exhaust gases escape (typically made from thermoset plastic); and adding a bulkhead closure that may contain a delay element and a cap to strongly secure the assembly in the casing and maintain the internal pressure necessary for operation. Manufacture and assembly of known castable composite rocket engines is therefore labor intensive, equipment limited, and difficult to automate.

BRIEF SUMMARY OF THE INVENTION

The present invention overcomes the deficiencies noted above for black powder and castable composite propellant compositions, and provides compositions with a functionally desirable specific impulse, a high combustion temperature, and improved manufacturing characteristics.

In one aspect, the invention relates to propellant compositions comprising a solid inorganic oxidizing agent; a nitrogen-containing fuel; and a burn rate catalyst.

The invention also relates to rocket engines (also referred to as rocket motors) packed with a composition comprising a solid inorganic perchlorate oxidizing agent, a nitrogen-containing fuel, and a burn rate catalyst.

In accordance with yet another aspect, the invention relates to high burn rate, high combustion temperature propellant compositions comprising a solid inorganic perchlorate oxidizing agent, a nitrogen-containing fuel consisting essentially of dicyandiamide; a burn rate catalyst which is preferably an oxide of copper, chromium, cobalt, manganese, iron, vanadium, or a mixture thereof; and a binder. The burn rate catalyst of these compositions may be particulate.

In accordance with still another aspect, the invention relates to an energetic composition comprising a solid inorganic oxidizing agent selected from the group consisting of potassium perchlorate and ammonium perchlorate; a nitrogen-containing fuel; and a burn rate catalyst, wherein said burn rate catalyst is nanoparticulate.

In accordance with one or more embodiments, the invention relates to a rocket engine comprising a case having a tensile strength of less than about 5,000 KPa and a composite propellant grain disposed in the case. The composite propellant grain can comprise an oxidizer, a burn rate catalyst, and a binder comprising a vinyl resin, and a polyfunctional ether.

In accordance with one or more embodiments, the invention relates to a method of preparing a pressable composite propellant formulation. The method comprises acts of mixing a vinyl resin, a vegetable oil, and a solvent to provide a binder mixture; adding a burn rate catalyst to the binder mixture; adding an amide to the binder mixture; and reducing the solvent from the binder mixture to about 7 wt % to provide a mixed formulation.

Described further herein are similar propellant compositions, as well as specifics regarding the relative amounts of the various components. In particular, the identities and relative amounts of oxidizing agent, nitrogen-containing fuel, burn rate catalyst, and binder in the compositions of the invention are disclosed. Also disclosed are methods for making and using the propellant compositions of the invention.

BRIEF DESCRIPTION OF THE DRAWINGS

The accompanying drawings are not intended to be drawn to scale. In the drawings, each identical or nearly identical component that is illustrated in various figures is represented by a like numeral. For purposes of clarity, not every component may be labeled in every drawing. In the drawings:

FIG. 1 is a graph showing the specific impulse of propellant compositions in accordance with one or more embodiments of the invention relative to black powder and typical castable composite formulations;

FIG. 2 is a graph showing the density impulse of propellant compositions in accordance with one or more embodiments of the invention relative to black powder and typical castable composite formulations;

FIG. 3 is a chart illustrating the preparation of propellant compositions in accordance with one or more embodiments of the invention;

FIG. 4 is a graph showing the baseline ballistic performance at various chamber pressures (psia) of the composite propellant in accordance with one or more embodiments of the invention wherein I_(d) refers to the Density Impulse, I_(sp) refers to the specific impulse (lb/sec), T refers to the combustion temperature (° F.), and MW refers to the exhaust gas molecular weight;

FIG. 5 is a graph showing the ballistic performance at various chamber pressures (psia) of black powder wherein Id refers to the Density Impulse, I_(sp) refers to the specific impulse (lb/sec), T refers to the combustion temperature (° F.), and MW refers to the exhaust gas molecular weight;

FIG. 6 is a graph showing the specific impulse performance of the composite propellant in accordance with one or more embodiments of the invention relative to the amount of potassium perchlorate oxidizer (KP) and fuel (DICY);

FIG. 7 is a graph showing the specific impulse performance, at various chamber pressures, of the composite propellant formulation in accordance with one or more embodiments of the invention at various binder concentrations;

FIG. 8 is a graph showing the burn rate at various chamber pressures of the composite propellant of the invention at various levels of SICOTRANS™ L 2715 D iron oxide;

FIG. 9 is a graph showing the ballistic performance at various chamber pressures (psia) of the composite propellant in accordance with one or more embodiments of the invention wherein I_(d) refers to the Density Impulse, I_(sp) refers to the specific impulse (lb/sec), T refers to the combustion temperature (° F.), and MW refers to the exhaust gas molecular weight;

FIG. 10 is a graph showing the ballistic performance at various chamber pressures (psia) of the composite propellant in accordance with one or more embodiments of the invention wherein I_(d) refers to the Density Impulse, I_(sp) refers to the specific impulse (lb/sec), T refers to the combustion temperature (° F.), and MW refers to the exhaust gas molecular weight; and

FIG. 11 is a graph showing the ballistic performance at various chamber pressures (psia) of the composite propellant in accordance with one or more embodiments of the invention wherein I_(d) refers to the Density Impulse, I_(sp) refers to the specific impulse (lb/sec), T refers to the combustion temperature (° F.), and MW refers to the exhaust gas molecular weight.

DEFINITIONS

As used herein, the term “plurality” refers to two or more items or components.

The terms “comprising,” “including,” “carrying,” “having,” “containing,” and “involving,” whether in the written description or the claims and the like, are open-ended terms, i.e., to mean “including but not limited to.” Thus, the use of such terms is meant to encompass the items listed thereafter, and equivalents thereof, as well as additional items. Only the transitional phrases “consisting of” and “consisting essentially of” are closed or semi-closed transitional phrases, respectively, with respect to the claims.

As used herein, the term “propellant composition” should be understood to encompass use of the claimed compositions in rocketry, pyrotechnics, military weapons and ammunition, as well as other applications, as propellant materials, as pyrotechnic materials, as explosive materials, as light generating materials, as heat generating materials, or as sound generating materials. As also used herein, the term “energetic composition” is intended to include the “propellant composition” defined above.

DETAILED DESCRIPTION OF THE INVENTION

This invention is not limited in its application to the details of construction, compositions, and/or arrangements of components or acts set forth in the following description. The invention is capable of being practiced and/or of being carried out in various ways or embodiments beyond those exemplarily presented herein and in the accompanying figures.

Energetic compositions have differing properties which are useful for specific purposes or applications. For example, some energetic compositions may be used for destruction, others for pushing or propelling, and still others for generating light or sound, and so on. Most, if not all, energetic compositions contain at least an oxidizer and a fuel, which may sometimes be in the same molecule. A binder may also be included to bind or keep the oxidizer and fuel together, as well as to prevent physical degradation and to increase mechanical strength. To specify or modulate the rate of oxidization of the fuel, and/or the compounded propellant's sensitivity to pressure or temperature changes, a burn-rate catalyst may also be added.

The individual components of the present invention were selected based on several factors including, inter alia, cost, availability, ease of use, and safety of handling. The propellant compositions of the invention, also referred to as VULCANITE™ EB-75 herein, are a significant improvement over both black powder and known castable composites, especially in model rocketry applications, because they can provide improved ballistic performance, improved volumetric and mass efficiency, a beneficial reduction of combustion products toxicity, and improvement in production efficiency, as described in more detail below. Indeed, propellant compositions in accordance with the present invention facilitate higher levels of automation; therefore, they allow greater production rates with a significant decrease in work-in-process inventories.

When the selected finely ground chemical components are mixed or compounded into an intimate mixture, the resulting propellant compositions of the invention may be granulated, pelletized, pressed or even “molded” via pressing or co-casting with other binders or propellants. This attribute makes the propellant compositions uniquely versatile for mass production with automated pressing equipment or other manufacturing methods deemed desirable. Because they can be processed in dry powder or granular form, the propellant compositions of the present invention are pressable, and are not subject to the viscosity, air entrapment, or pot-life problems inherent to castable propellants. As such, rocket engine production is easily automated by means of single or multiple feed, hydraulic pressing machinery.

Particularly regarding its use in manufacturing of rocket engines, the propellant compositions of the invention can be used as granules or pellets which can then be pressed into the engine casing, or they can be pressed into molds to form “grains,” the mass of solid propellant used in rocket engines or other applications. Grains can then be inserted into engine casings or sold separately, allowing the user to load the casing on an “as needed” basis.

Grains or pellets of the compositions of the invention have improved ballistic performance as discussed below, and will find utility in a variety of other applications and devices, such as, but not limited to, ignition transfer pellets used in piccolo tubes, pyrotechnics, artillery shell igniters, hand signals, maritime smokes and signals, expelling charges and various other applications.

The propellant compositions of the invention may be used as a high burn rate (at least about 5/10 inch per second at a chamber pressure about 150 pounds per square inch absolute (psia)) propellant for rocket engines, including, for example, in model rocket engines, because it allows easily producible grain geometries and operation at reduced chamber pressures. The solid nature of the propellant compositions of the invention allows grains of various geometries to be produced using automated procedures such as molding and pressing. Because the propellant composition can be tightly packed or pressed to a uniform density, it burns evenly. Pressed grains are typically of higher uniform quality because consolidation under pressure prevents formation of air bubbles or cavities, as can occur in liquid mixtures. Such cavities are undesirable because they result in burn rate variability. Depending on the grain geometry selected, the operating chamber pressure can be modified, as is known in the art.

The propellant compositions of the invention also have a high combustion temperature (as high as about 4,000° F. and even greater). An advantageous consequence of this high combustion temperature is that the composition is more efficient because the higher temperature imparts greater energy to the propulsive gases. As a result, less composition is required to produce the same total impulse as a given weight of black powder, as shown in Table 2, where the Isp of the propellant composition of the present invention can be about 1.73 times that of black powder and can produce a combustion temperature nearly 1.5 times that of black powder. Generally, the combustion temperature is at least about 3,500° F., and in some cases, the combustion temperature is greater than about 4,000° F. Less propellant mass is required to produce the same total impulse as a given weight of black powder; therefore, it provides an increase in volumetric efficiency at comparable densities.

In addition, the propellant compositions of the invention have improved ballistic properties. Compositions of the present invention have a specific impulse of at least about 100 to about 120 seconds at about 100 psia (pounds per square inch absolute), and thus are more energetic than black powder, which typically has a specific impulse in a range of 80 to 101 at a pressure of about 100 psia. In selected embodiments, the propellant compositions of the invention have a specific impulse of at least about 120% that of black powder. See Table 2, below.

Specific impulse, Isp, may be calculated according to the following formula: ${Isp} = {\frac{1}{g}\sqrt{\left( \frac{2\gamma}{\gamma - 1} \right){\left( \frac{{RT}_{c}}{M} \right)\left\lbrack {1 - \left( \frac{P_{e}}{P_{c}} \right)^{\frac{({\gamma - 1})}{\gamma}}} \right\rbrack}}}$ where

-   -   γ is the ratio of the specific heats of the combustion gases         (Cp/Cv),     -   R is the universal gas constant,     -   T_(c) is the absolute combustion temperature,     -   g is the gravitational acceleration constant,     -   M is the average molecular weight of the exhaust gases,     -   P_(c) is the combustion chamber pressure, and     -   P_(e) is the exhaust gas pressure at the nozzle exit.

To maximize specific impulse, the molecular weight of the exhaust gas and the relative amounts of solid combustion byproducts should be minimized, and the combustion temperature and ratio of chamber pressure to exit pressure should be maximized. In rocketry applications, especially in model rocketry cases, the ratio of chamber pressure to exit pressure is typically determined by the design of the rocket engine itself in conjunction with the ballistic properties of the propellant.

Another benefit of the propellant composition of the invention is the ease of obtaining reproducible performance. The chemicals used are consistent in composition and purity and are consistent in the amount of energy delivered per unit of material. Additionally, the components of the compositions are less subject to change from the absorption of moisture than, for example, the components of black powder. These features represent a further advantage over conventional composite propellant compositions and black powder.

In comparison, black powder performance (burn rate, energy produced, burn temperature, etc.) can vary significantly from batch to batch. Primarily, this is because black powder is made with charcoal. The properties of charcoal, which is manufactured from wood, can vary greatly depending on the species of wood, where the wood was harvested, climate conditions during the life of the wood, and the temperature and manner in which the wood is converted into charcoal. Furthermore, charcoal tends to absorb moisture from the atmosphere. The use of black powder in the manufacturing of rocket engines requires extensive batch testing of the powder prior to production so that its ballistic characteristics are known. The compositions of the present invention require less extensive pre-manufacturing testing because the chemical components do not vary. Consequently, the compositions of the present invention do not exhibit the same batch to batch variability.

The compositions of the invention produce exhaust gases (combustion gases) of low average molecular weight, less than about 45, and preferably about 39 to about 40. By comparison, the average molecular weight of exhaust gases from black powder are higher, about 48, and exhaust gases from a typical castable composite have an average molecular weight of about 23. Lower average molecular weight combustion products are easier to accelerate to higher velocities, making the compositions of the present invention more efficient propellants than black powder. Additionally, in some rocketry applications, the invention provides less solid residue, which can alter the performance of a rocket engine. Further, reducing the solid residue also is desirable because it produces less build-up on launch equipment.

The very low percentage of non-expandable solid by-products in the exhaust stream, preferably less than about 5 mole %, increases efficiency. By comparison, black powder engines typically have 15.96 mole % solid by-products in the exhaust stream. Castable compositions typically have about 1 mole % solid by-products. This makes the propellant compositions of the present invention more efficient than black powder in creating energy, with less undesirable residue. Furthermore, fewer particulates in the hot exhaust stream reduce the risk of fire to the surroundings in a variety of applications.

The propellant compositions of the invention can also have a high burn rate coefficient as compared to typical castable compositions. The burn rate, r_(b), may be expressed as r _(b) =a(P _(c))^(n),

wherein a is the burn rate coefficient, P_(c) is the chamber pressure, and n is the burn rate exponent. As shown in Table 1, below, in comparison with typical castable compositions, the propellant compositions of the invention combine a reasonably low burn rate exponent with a relatively high burn rate coefficient, as obtained from Crawford Bomb and Micro Motor firings, as is known in the art. This combination of relatively low burn rate exponent, less than about 0.5, with relatively high burn rate coefficient, greater than about 0.1, is both unique and beneficial. It is beneficial because, inter alia, it allows for higher mass flow rates at low chamber pressures in a range of about 50 to about 200 psia, pressures at which known castable compositions do not function reliably. TABLE 1 Comparative Burn Rates Burn Rate (above Burn Rate Burn Rate about 100 psia) Exponent Coefficient Black Powder r = 0.5974P_(c) ^(0.0789) 0.05-0.11 0.50-0.70 Typical Castable r = 0.0482P_(c) ^(0.3607) 0.32-0.70 0.02-0.06 Composition VULCANITE ™ EB-75 r = 0.1033P_(c) ^(0.3613) 0.36-0.50 > about 0.10 Propellant Composition

The low burn rate exponent reduces variation in performance based on pressure. A low burn rate exponent greatly reduces the pressure sensitivity of the propellant, thereby allowing larger burning surface area changes without extreme pressure increases. This is beneficial for use with the materials typically used in the construction of rocket engines, including, for example, model rocket engines, where the inner diameter of the casing may vary from casing to casing and even within individual casings. The lower sensitivity can be especially significant at low combustion pressures because any variation at such pressures can attenuate any undesirable variations in burn rates. Thus, in embodiments of the invention wherein low-strength casings are utilized, burn rate exponent sensitivity can lead to unpredictable performance characteristics. The burn rate exponent of compositions of the invention may be less than about 0.5.

In accordance with some embodiments, the invention relates to propellant compositions comprising a solid inorganic perchlorate oxidizing agent; a nitrogen-containing fuel; and a burn rate catalyst. The burn rate catalyst can be an oxide of copper, chromium, cobalt, manganese, iron, vanadium, or a mixture thereof. The burn rate catalyst is preferably a high surface area particulate. Other embodiments of the invention pertain to rocket engines packed with propellant compositions comprising a solid inorganic perchlorate oxidizing agent, a nitrogen-containing fuel, and a burn rate catalyst. In accordance with some embodiments, the invention relates to high burn rate, high combustion temperature propellant compositions comprising a solid inorganic perchlorate oxidizing agent, a nitrogen-containing fuel consisting essentially of dicyandiamide, a burn rate catalyst, and a combustible binder. An advantageous example of the propellant compositions of the invention is a high burn rate, high combustion temperature composition comprising about 64 to about 72 wt % of a solid inorganic perchlorate; about 15 to about 23 wt % of a nitrogen-containing fuel; about 0.5 to about 10 wt % of an oxide of one or more of copper, chromium, cobalt, manganese, iron, vanadium; and about 0.75 to about 12 wt % of a combustible binder.

The burn rate catalyst can be instrumental in lowering the burn rate exponent of the propellant compositions of the invention. The catalyst is preferably a high surface area particulate, and is typically irregularly-shaped. The burn rate catalyst typically has a large surface area, which can increase the catalytic activity thereof. The burn rate catalyst can comprise one or more inorganic oxides of a metal. In accordance with some embodiments of the invention, the average size of the catalyst is relatively small and may thus be characterized as nanoparticulate. Thus, for example, the metal oxides can be utilized in the compositions of the invention as nanoparticulates wherein at least one dimension thereof is in the nanoscale domain. The nanoscale dimension can be a smallest dimension or a largest dimension of the catalytic particle. Further embodiments of the invention contemplate the use of a mixture of catalysts having a variety of shapes. For example, some embodiments of the invention can include a burn rate catalyst having an acicular shape in combination with one or more burn rate catalysts having a globular shapes.

The propellant composition of the present invention can comprise one or more thermally conductive species. In accordance with some aspects pertinent to one or more compositional embodiments of the invention, a particulate ingredient thereof can provide or at least facilitate heat transfer during the combustion. For example, the thermally conductive species can have a high shape aspect ratio such that the species can be characterized as being acicular. Such habits can facilitate thermal conductivity through the propellant composition body, e.g. the grain, because a long dimension thereof can thermally expose a region of the propellant composition to a higher temperature when the bulk propellant composition has a lower thermal conductivity, relative to the thermally conductive material. Thus, the burn rate catalyst, in accordance with some embodiments of the invention, can facilitate desirable propellant characteristics by chemically and thermally modifying the behavior thereof.

The surface area of the catalyst is preferably greater than about 50 m²/g. The preferred catalysts are metal oxides that can modify, e.g., accelerate, the reduction/oxidation (redox) reactions associated with propellant combustion. An example of a burn rate catalyst suitable in the compositions of the invention include those commercially available as SICOTRANS® Red L 2715 D iron oxide, available from BASF Corporation, Florham Park, N.J. In some compositions of the invention, the burn rate catalyst may constitute about 0.1 wt % to about 15 wt % of the composition and, in some cases, may be about 0.25 wt % to about 10 wt %, based on 100 wt % of the propellant weight.

A binder may also be used in the propellant compositions of the invention. The binder can comprise at least one resin material. The binder can further comprise at least one material that comprises at least one functional group that can react with one or more components of the binder. For example, the binder can comprise a compound having a functional group that reacts with the polymeric precursor species and/or the at least one modifier. Indeed, the binder materials can be selected to provide or facilitate the transformation of propellant composition properties from a viscous material, capable of being disposed into a case, to a dimensionally stable solid propellant grain. The binder materials can be selected to provide any further desirable propellant grain properties. For example, the binder precursor materials can be selected to provide a propellant grain having high burn rates at low chamber pressures. Moreover, the species comprising the propellant compositions of the invention can also be selected to provide desirable energetically stable compositions. The binder of one or more propellant compositions of the invention can provide a matrix that can serve as carrier for the oxidizer and/or a fuel. In accordance with some embodiments of the invention, the binder matrix can also provide or serve as a fuel, or even as a fuel supplement, during combustion. Typically, the binder matrix can also facilitate processing of the composite propellant composition. Indeed, in some embodiments of the invention, the binder matrix serves as a carrier for at least one oxidizer and/or at least one burn rate catalyst such that the composite can be pressed into a case or casing and/or form propellant grains therein. In accordance with some aspects pertinent to one or more embodiments of the invention, the binder matrix can facilitate production of propellant grains by providing compositions with desirable rheological properties during preparation and processing operations thereof and further provide propellant grains that have desirable mechanical properties during, for example, storage, transport, and/or uses thereof. Thus, one or more propellant compositions of the invention can have desirable rheological properties during, for example, mixing and/or pressing, and can further have desirable, stable mechanical properties during propulsive service. For example, one or more components of the binder matrix can reduce the viscosity of the mixture to facilitate mixing.

The binder may prevent the propellant composition granules from being degraded, e.g., broken down into potentially dangerous dust during the manufacturing processes and/or during transport and storage thereof. The binder may also increase the mechanical strength of the composition, which, in some embodiments of the invention, can be effected by chemical cross-linking. For example, one or more ingredients or species comprising the binder can react to form a cross-linked network, thereby imparting greater mechanical strength to the propellant composition structure. The binder may include one or more modifiers or be modified by an ingredient to provide desirable processing characteristics. For example, the binder may comprise one or more plasticizing modifiers. The preferred binders are understood to assist, rather than interfere with, the propellant functionality. Various ingredients can be utilized in the binder matrix. For example, the binder matrix can comprise a polymeric material or a precursor thereof that forms or reacts to provide a dimensionally stable propellant grain. Preferred binders and/or components thereof are combustible and produce low molecular weight by-products, and are typically combustible organic polymers. Unlike inorganic binders such as silica, the preferred binders do not increase the quantity of non-expandable solid combustion products. In some cases, the binder can comprise one or more modifiers that can modify a Theological property of a mixture thereof, e.g., reduce a viscosity of the precursor mixture. Thus, in accordance with some aspects of the invention, the modifier can serve, among other things, as a plasticizer of the propellant composition to provide desired rheological properties.

In accordance with one or more embodiments of the invention, the binder can comprise at least one resin, at least one modifier, and/or at least one reactive agent. One or more components comprising the binder can include chemically reactive species. In some cases, one or more components of the binder can include functional groups, which can react with other functional groups pendant on or a part of other components of the propellant composition. Non-limiting examples of compounds, such as resins, which can be utilized in the propellant compositions of the invention, include those comprising one or more vinyl functional groups and/or those comprising one or more hydroxyl functional groups. Non-limiting examples of compounds, such as modifiers, which can be utilized in the propellant compositions of the invention, include functionalized species, e.g. having mono or polyfunctional groups like polyfunctional ethers. Specific examples of such polyfunctional ethers include di- and trigycidyl ethers or epoxidized organic compounds.

The propellant compositions of the invention may be 0.5 wt % to about 15 wt %, or more particularly 0.75 wt % to about 12 wt % binder. The binder may be comprised of at least one organic polymer alone or in combination with a plasticizer such as dioctyl adipate, dioctyl sebacate, hydrocarbon ester tackifier, and combinations thereof.

Materials of the binder can be selected to have a high density, which can provide desirable ballistic performance. The selected binder materials can also provide mechanically stable grains, especially when loaded into paper motor cases. For example, the binder materials can provide grains that can be readily pressed into cellulose-based cases without crumbling to excessive dust. In some cases, the binder materials can effect moldability of the composite propellant mixture in the case. Such materials can also accommodate automated pressing operations, e.g., by hydraulic pressing. Examples of preferred binder materials include, but are not limited to, those comprising alkyl acetate, vinyl acetate, and/or vinyl chloride resins. Further, the binder materials can include one or more plasticizing agents, tackifying agents, bonding agents, and/or wetting agents. Such agents can also be reactive or have one or more reactive functional groups. As discussed herein, the functional group can react with itself and/or one or more other components in the composition. The desirable binder materials can provide combustion products having a low average molecular weight and high combustion temperatures, especially with respect to black powder. For example, some amides have a typical combustion temperature of greater than about 4438° F., which is relatively higher than the combustion temperature of black powder (about 3426° F.). Further, the binder materials can be selected to provide gaseous combustion products, especially with the disclosed oxidizing agents discussed herein. The binder materials can also be selected based on toxicity such that preferred candidates are considered to be non-toxic, as defined or regulated by, for example, governmental entities.

Non-limiting examples of components of binders effective in the practice of the present invention include polyvinylchloride, polyvinylacetate, polyvinylalcohol and/or copolymers thereof such as the solution vinyl resins commercially available as UCAR™ VAGH, VAGD, VAGC, and VROH resin, each of which is available from The Dow Chemical Corporation, Midland, Mich., GEON-121 (available from B. F. Goodrich Corp.), poly (2-ethyl-2-oxazoline), and epoxy or acrylate resin, epoxidized trimethylolpropane, trimethylol ethane triglycidyl ether, such as HELOXY™ Modifier-44 from Shell Chemical Company or Resolution Performance Products, Houston, Tex., epoxidized soybean oil, and combinations or mixtures thereof. Non-limiting examples of agents include flexibilizers or modifiers such as those commercially available as HELOXY™ esters, e.g., HELOXY Modifier 505 castor oil polyglycidyl ether, from Resolution Performance Products, Houston, Tex.

Non-limiting examples of compounds that can react with one or more of the resin and the modifier include those that are nitrogen-containing species, thus serving, at least partially, as a fuel source. The fuel component of the propellant composition typically reacts with the oxidizer to produce the propulsive exhaust gas. In some cases, the fuel species can react with a functional group of a component of the propellant composition and provide the propellant composition with desirable mechanical, e.g., rheological properties, and also react during propulsive service with the oxidizer. The nitrogen-containing fuel is generally a cyano-, amide-, or amine-containing material, or a mixture thereof. Some examples include acrylonitrile, amino tetrazole, aminoguanidinium bitetrazole, ammonium dicyanamide, bistriaminoguanidiniumdecacarborane, bis(trinitroethyl)nitramine, calcium bitetrazole, dicyandiamide (or cyanoguanidine), nitroaminoguanidine, triaminoguanidine, and triaminoguanidinedicyanamide. The fuel component of the propellant compositions of the invention may be about 10-30 wt %, and more particularly, about 15-23 wt %.

An example of a composition according to the invention consists of about 62-72 wt % potassium perchlorate, about 15-30 wt % dicyandiamide, about 2-10 wt % iron oxide, and about 3-12 wt % of a VROH base stock comprising a plasticizer and VROH mixed in a solvent and subsequently evaporated. The VROH base stock mixture comprises about 3:2 VROH to plasticizer (e.g., dioctyl sebacate). More particularly the composition may consist of about 68 wt % potassium perchlorate, about 19 wt % dicyandiamide, about 5 wt % iron oxide, and about 7.5 wt % of a VROH base stock comprising a mixture of about 58.8:41.2 VROH to dioctyl sebacate.

The oxidizer typically facilitates combustion, such as by oxidation, of one or more of the ingredients of the propellant composition. The oxidizer can comprise an oxygen-containing compound. Typically, the oxidizer can comprise an ammonium, alkali, or alkali metal salt of an oxygen-containing compound. In accordance with one or more embodiments of the invention, the oxidizer facilitates stable combustion or deflagration at low pressures, typically at less about 200 psia. The oxidizing compound can include those considered as suitable for high pressure conditions and/or associated with high flame temperatures or high burn rates. The oxidizer can also be selected based on material density. Typically, oxidizing agents having a high density are suitable in some formulations of the invention. Further, the selected oxidizer is preferably non-hygroscopic to facilitate storage and ease of use. For example, in accordance with one or more embodiments of the invention, the oxidizer can comprise an ammonium or alkali perchlorate. However, in some cases, the oxidizer is selected to provide compositions having high burn rate exponents and poor low-temperature ballistic properties. Although such properties are typically undesirable in certain applications of the invention, they can be addressed by the utilization of other ingredients, such as the burn rate catalyst. The solid inorganic oxidizing agent may be potassium perchlorate or ammonium perchlorate. In some embodiments, the oxidizing agent is about 60 to about 75 wt %, or more particularly, about 64 to about 72 wt % of the composition.

The propellant compositions of the invention can be prepared by any suitable technique. FIG. 3 presents a flow chart exemplarily illustrating acts that may be utilized to prepare propellant compositions in accordance with one or more embodiments of the invention. Thus, any suitable equipment may be utilized to prepare the grain forming compositions of the invention. Further, the acts of propellant composition preparation may be performed in any suitable order or sequence. Indeed, in accordance with one or more embodiments of the invention, the propellant compositions of the invention can be prepared by utilizing pre-batching techniques which typically involves preparing an aggregate of two or more components of the composition. Further, where desired, or in some cases, necessary, preparatory procedures can be performed to render any ingredients suitable or active in the propellant composition.

For example, a pre-batch or pre-blend can be prepared by mixing one or more vinyl resins and one or more solvents, and, optionally, one or more modifiers. Pretreatment or preparatory procedures, such as drying, can be performed prior to or after the pre-batch preparation. The one or more solvents can be any suitable solvent that is selected to dissolve the resin. Examples of suitable solvents include those that readily evaporate such as esters, including, but not limited to, ethyl acetate. Any suitable amount of solvent can be used. Typically the amount of solvent utilized provides desirable pre-blend mixture rheological properties. For example, the amount of ethyl acetate can be about equal to the amount, e.g., the weight, of vinyl resin. In accordance with some embodiments of the invention, the amount of solvent utilized can be selected to provide a target pre-batch viscosity while resulting a minimal amount, e.g., about 7 wt %, in the aggregate propellant composition or mixed formulation.

The pre-batch mixture can then be charged into a mixing apparatus wherein additional components or ingredients can be added and mixed. For example, dried burn rate catalyst can be mixed with the pre-batch in a vertical mixing apparatus until the burn rate catalyst, e.g., acicularly-shaped, nanoparticulate iron oxide, has been wetted by the pre-batch mixture. Drying of the catalyst can performed at, for example, a temperature of about 350° F. for about three hours under a vacuum of about 29 inches of mercury to achieve a moisture content close to about 0%. Examples of suitable mixing apparatus include those commercially available from Charles Ross & Son Company, Hauppauge, N.Y. Mixing can be performed under vacuum. Any suitable vacuum level can be utilized. For example, a vacuum of about 29 inches of mercury can be applied during this mixing operation.

The oxidizer can then be added. One or more preparatory procedures can likewise be performed on the oxidizer. For example, potassium perchlorate can be ground to have a desired particulate size, e.g., about 10 μm, in a hammer mill. Further the oxidizer can also be dried in substantially the same fashion as described with respect to the burn rate catalyst.

One or more binder material-reactive components can be added. For example, dicyandiamide can be added. As with the previously-mixed ingredients, preparatory procedures such as grinding and/or drying may be performed prior to addition thereof in the mixing apparatus.

Mixing to incorporate any of the propellant ingredients can be performed for any suitable period and under any suitable vacuum condition. Moreover, mixing can be performed at any suitable temperature but is typically performed at about room temperature.

Following addition of all ingredients, mixing can be maintained under an applied vacuum, e.g., about 29 inches of mercury, until the mixture provides desired rheological properties. For example, vacuum mixing can be performed until about 7 wt % of solvent remains.

The mixed composition can also be further processed to remove any entrapped gas. The mixed composition can also be rendered in any desired shaped. For example, the composition can be extruded to remove trapped gas and then form pellets. The extruded pieces can have any suitable or desired size. For example, the pieces can be extruded through a die having an aperture diameter in range of about 0.06 to about 0.08 inch. Further, the extruded pieces can have any desired aspect ratio. For example, the length to diameter can be about 2:1 to about 3:1.

Further processing can optionally be performed on the extruded pieces prior to disposing into a case, which is further described below. For example, the extruded pieces can be shaped or formed to be spherical in a spheronizer. In some cases, further drying can be performed in, for example, a fluidized bed to remove any residual solvent. Any suitable fluidizing gas can be used at any suitable temperature that facilitates solvent evaporation. For example, an inert gas at a temperature of about 120° F. can facilitate solvent evaporation.

Further modifications directed to the propellant formulation of the invention include the addition of agents that provide or modify the other physical or performance properties. For example, one or more glazing or anti-caking agents may be utilized to reduce the likelihood of clumping or caking of any of the ingredients. Examples of such agents include, but are not limited to, stearic acid coatings. Further, agents that preserve one or more of the propellant formulation ingredients may be utilized. For example, a sterically-hindered amine anti-oxidant may be utilized to facilitate stable long term storage of the vinyl resin, the functionalized modifier, a mixture thereof, and/or the resultant composite propellant. Other agents such as one or more sterically-hindered phenol compounds, which may also be utilized with amide-based sequestering or chelating agents, may facilitate processing by at least partially inhibiting any reaction between, for example, the functionalized fuel and the functionalized modifier components. Such agents would thus allow for longer work-in-process periods. Further, anti-static agents may be utilized to reduce the electrostatic sensitivity and accommodate air conveyor-based transport systems and techniques.

The present invention also relates to a method of manufacturing a rocket engine in which propellant compositions of the invention are loaded into a rocket engine chamber. Such a method can be used to make one engine at a time using a hand-operated pressing machine or used with automated equipment capable of making thousands of engines per day.

Processed in the dry powder or granular form, the pressable propellant composition of the present invention is typically not subject to the viscosity, air entrapment, or pot-life problems inherent to castable propellants. As such, pressable propellant rocket engine production is easily automated by means of multiple feed, hydraulic pressing. In comparison to castable motor production, this method produces little waste material, can be quality inspected on a near real time basis and results in a minimum of product held as “work in progress.” Further, the composite propellant compositions in accordance with some embodiments of the invention advantageously undergo a B-stage cure, which can be at ambient conditions, which in turn can advantageously increase process equipment availability. After pressing, a first batch of propellant grains can be removed from the pressing tools and processing a next batch can be initiated. Because the propellant compositions are pressable and generally retain a shape or configuration after pressing, the first batch need not be fully cured before removal from the pressing equipment. The availability and processing capacity is thus increased, especially with respect to castable compositions.

Because the propellant composition of the present invention contains little or no moisture sensitive raw materials and is post processed in dry powder or granular form; it is typically not subject to the same processing constraints as castable compositions. Processing is straightforward and very flexible with respect to suitable mixing equipment.

In another embodiment of the invention, rocket engines are manufactured by pressing clay or other suitable heat resistant material into a casing to form the rocket engine nozzle. Alternatively, the nozzle may be pre-formed and then inserted into the casing. The casing is preferably a multi-layered paper or cardboard in any suitable configuration, such as corrugated or laminated or combinations thereof; may be made from plastic, fiberglass, a filament-wound glass/epoxy composite, paper phenolic, plastic phenolic, and aluminum or its alloys, as is known in the art.

The case material is selected based on several factors including, but not limited to, cost, availability, and mechanical properties. As noted, typical cases can comprise paper, cellulose-based materials, or non-metallic materials. Thus, the specific gravity of the material comprising the case can be less than about 2. The case can further comprise a composite or assembly of various types of materials to provide desired mechanical properties. In accordance with some embodiments of the invention, the case is constructed and arranged based on a chamber operating pressure of less than about 300 psia, and in some cases, based on a chamber pressure of about 100 psia. Thus, in some cases, the case have a burst strength of less than about 2,000 KPa, and/or a tensile strength of less than or equal to about 5,000 KPa. However, where it is advantageous to do so, the burst strength can be about 1,000 KPa, or less. For example, where a reduced cost can be realized without a sacrifice or an increase in the likelihood of failure, the case can be constructed and arranged utilizing a material that has a burst strength of about 1,000 KPa, or less.

The propellant composition powder (for producing thrust) is fed into the casing in incremental amounts sufficient to achieve a uniform pressed density, and compressed at high pressure (about 10,000 psia or higher) to form a single propellant grain with uniform density. Alternatively, the grain may be pre-formed outside of the casing, and then inserted into the casing. The required amount of propellant composition needed to produce the desired total impulse is calculated by dividing the desired total impulse by the specific impulse of the propellant compound, as is known in the art. For example, a type “C” engine has a desired total impulse of 10.0N-S therefore, 7.5 grams of the propellant composition is used.

Delay powder is then fed into the engine casing and pressed to achieve the desired time delay prior to igniting the ejection charge. Alternatively, it can be formed into a unit and inserted in the casing after forming. The amount of delay powder used may be calculated based on, inter alia, the burn rate of the delay powder, the size, e.g., diameter of the engine and the desired time requirement.

Ejection powder, typically black powder, is then put into the engine casing. The purpose of the ejection powder is to provide gas to deploy the recovery mechanism.

Clay or similar heat resistant material is then inserted into the casing and pressed at a low pressure sufficient to retain the ejection powder in the casing yet allow the release of the ejection gases in order to activate the recovery mechanism.

This invention is further illustrated by the following examples, which should not be construed as limiting.

EXAMPLES

The function and advantages of these and other embodiments of the invention can be further understood from the examples below, which illustrate the benefits and/or advantages of the one or more systems and techniques of the invention but do not exemplify the full scope of the invention. In the examples, the performance of the various propellant formulations was characterized utilizing, inter alia, a widely available program, PROPEP. Because the associated JANAF material database thereof did not include VROH resin and HELOXY™ 505, thermochemically similar species, GEON 121 and dioctyl adipate, were respectively substitituted therefor.

Example 1 Preparing a Propellant Composition of the Invention (VULCANITE™ EB-75 Composition A)

The mixer is charged with 100% of the VROH binder (Dow Chemical) in solvent solution form. Under mixing action in a planetary mixer, the prescribed amount of mono-modal particle size potassium perchlorate below is added to the VROH solution until a uniform, paste like consistency is obtained. Other types of suitable mixers are known in the art, such as whip mixers, twin screw mixers and Mueller mixers. The dicyandiamide and iron oxide are then added to the mixer in respective order and processed until a uniform, paste-like consistency is achieved. In one-kilogram batches such as described here, this is a matter of minutes. The entire mixing process is achieved easily within a very short period of time, usually in 15 minutes or less.

-   -   0.92% VROH hydroxyl-modified vinyl copolymer (The Dow Chemical         Company, Midland Mich.)     -   64.22% potassium perchlorate (Service Chemical Inc., Hatfield,         Pa.)     -   27.52% dicyandiamide (Air Products and Chemicals, Inc.,         Allentown, Pa.)     -   7.34% iron oxide nanoparticles (SICOTRANS® Red L 2715 D)

Once a uniform mixture is obtained, again, a matter of minutes, the material is removed from the mixer and excess solvent removed by evaporation If evaporation is accomplished manually, this will take several hours. If a dryer, such as a hot air recirculating tunnel is used, this will take minutes. Upon reaching a predetermined solvent content whereby the mixture is a pliable mass capable of being screened without being so fluid as to stick together, or too dry to be screened, the material is screened to a desired particle size. For this example, the particle size was an 8-mesh market grade; the particle size can be varied according to the application. The particles are then allowed to dry completely over a period of hours, if accomplished manually, prior to packaging.

Example 2 Preparing Another Propellant Composition of the Invention (VULCANITE™ EB-75 Composition B)

Following the steps described above, a propellant composition B was manufactured using 68.04% potassium perchlorate, 19.22% dicyandiamide (≧7μ), 5.66% iron oxide nanoparticles, and 7.08% DOS modified VROH. The DOS modified VROH is prepared by mixing 59.52% methyl ethyl ketone, 23.81% VROH, and 16.67% dioctyl sebacate under agitation until dissolved.

Example 3 Engine Manufacture Using a VULCANITE™ EB-75 Propellant Composition of the Invention

Clay (about 2.5 grams) is fed and pressed into the multi-layered paper casing for a rocket engine to form the rocket engine nozzle. The propellant composition powder is then fed and pressed into the engine casing in increments sufficient to achieve a uniform pressed density, at a high pressure of about 10,000 psia, to form a single propellant grain of uniform density. The amount of powder used depends upon the type of engine being manufactured and is determined by calculating the amount needed to produce the desired total impulse. In this example, a type “C” engine was made, the total impulse for which is no more than 10.00 N-S. Therefore, about 7.5 grams of propellant composition was used.

Delay powder is then fed into the engine casing and pressed to achieve the desired time delay prior to igniting the ejection charge. In this example, a time delay of less than 8 seconds was desired, and 1.5 grams delay powder was added, based on internal casing diameter and burn rate calculations. Delay powder used in this example was PYRODEX® HF-20™ from Hodgdon Powder Company, Inc., Shawnee Mission, Kans., but may be selected from other slow burning compositions (about 0.05 inches per second) as is known in the art.

Ejection powder (black powder) is then put into the engine casing to create gas to deploy the recovery mechanism. The amount of ejection powder used is determined by engine casing size and estimated size of rocket to be flown, and ranges from about 0.5 grams for a “C” type engine to about 1.2 grams for larger engines.

Clay, about 0.5 grams for a “C” engine to about 1.2 grams for larger engines is then inserted into the casing and pressed at a low pressure to retain the ejection powder in the casing but still allow the release of the ejection gases in order to activate the recovery mechanism.

During and after manufacture, sample engines are periodically tested to ensure that they function as expected and otherwise continue to meet engine performance specifications. After “aging” for at least ten days, engines may be retested to ensure that they continue to meet specifications.

Example 4 Comparison of Specific Impulse (I_(sp)) of Black Powder and Compositions of the Invention

The combustion characteristics of black powder and a propellant composition of Example 6 were compared based on theoretical values. TABLE 2 VULCANITE ™ EB-75 Black Powder Composition C Combustion temperature 2221 4313 (° F.) Exhaust gas pressure at nozzle exit 12.1 14.7 (psia) Combustion chamber pressure 100 50 (psia) Molecular weight of exhaust 48.341 39.159 mixture Total Exhaust solids 15.96 2.61 (mole %) Cp/Cv 1.1318 1.2059 Delivered I_(sp) 75-80 148.6 (lb. seconds/lb.)

As can be seen from the table, the propellant composition C of the invention (Table 3, infra) showed a Δ I_(sp)=80.13% increase in specific impulse (I_(sp)) over black powder. This was accompanied by a Δ solids of −83.65% reduction in total exhaust solids, and a ΔT 94.19% increase in combustion chamber temperature.

Example 5 Combustion and Safety Characteristics

The granular propellant composition VULCANITE™ EB-75 Composition B of Example 2, consisting of 68.04% potassium perchlorate, 19.22% dicyandiamide (≧7μ), 5.66% iron oxide nanoparticles, and 7.08% DOS modified VROH (prepared from 59.52% methyl ethyl ketone, 23.81% VROH, and 16.67% dioctyl sebacate) was found to have the following characteristics:

-   -   Auto-ignition temperature: greater than about 500° F.     -   r=0.1033 P_(c) ^(0.3613).     -   Failed to explode when subjected to DOT Impact Test.     -   Only 0.067% weight loss per 200 g at 75° C. after 48 hours.

The flame temperature of the composition was calculated to be 4250° F., the molecular weight of the exhaust mixture was 38.5, and C_(p)/C_(v) was approximately 1.2.

Example 6 Calculated Thermodynamic Data for Various Compositions of the Invention

Using a Propellant Evaluation Program (PROPEP) widely used and available via the World Wide Web from, inter alia, The Gas Dynamics Lab, thermodynamic equilibrium data for various propellant compositions of the invention was calculated. The results appear in Table 3, below. Table 4 lists the composition of the propellant formulations evaluated in Table 3. TABLE 3 VULCANITE ™ EB-75 Composition Composition C Composition D E Combustion Temperature 4313 4091 4365 (° F.) Exhaust gas pressure at 14.7 14.7 14.7 nozzle exit (psia) Combustion chamber 50 50 50 pressure (psia) Avg. molecular weight of 39.159 28.90 33.466 exhaust products C_(p)/C_(v) 1.2059 1.2383 1.2275 I_(sp) 148.6 166.5 160.5 (lb. seconds/lb.)

TABLE 4 VULCANITE ™ EB-75 Composition C Composition D Composition E Potassium 68.04% 56.04% 60.04% perchlorate Nitrogen- 19.22% 33.58% 29.58% containing dicyandiamide triaminoguanidine triaminoguanidine species dicyanamide iron oxide 5.66% 5.66% 5.66% GEON-121 4.16% 2.78% 2.78% dioctyl adipate 2.92% 1.94% 1.94%

Example 7 Comparison of Specific Impulse and Density Impulse

The graphs in FIGS. 1 and 2 plot the specific impulse and density impulse of Propellant Compositions C to E, above, relative to black powder and a typical castable propellant. As can be seen, there is a significant increase in mass and volumetric efficiency of the compositions of the present invention over black powder.

Example 8 Propellant Composition

A propellant composition, indicated as “Baseline Formulation,” in accordance with one or more embodiments of the invention was prepared from a base stock mixture or pre-blend comprising the ingredients listed in Table 5 by mixing the ingredients listed. The pre-blend mixture was then charged in a vertical mixer. To the pre-blend mixture, a burn rate catalyst, an oxidizing agent, and a fuel source were added in the amounts listed in Table 6. Preparation of the propellant composition was performed in substantial accordance with the flow chart presented as FIG. 3.

The ballistic performance, at various chamber pressures, of the composite propellant composition is presented in FIG. 4. The ballistic performance was simulated by utilizing PROPEP code (June 1998 version), substituting GEON-121 for VROH, and dioctyl adipate for HELOXY™ 505 modifier. TABLE 5 Pre-blend Mixture Concentration Ingredient (wt %) Resin 23.81 (VROH) Modifier 16.67 (HELOXY ™ 505 Modifier) Solvent 59.52 (Methyl Ethyl Ketone)

TABLE 6 Baseline Formulation Concentration Ingredient (wt %) Potassium perchlorate 68.04 (particle size of about 10 μm) Dicyandiamide 19.22 Iron Oxide 5.66 (SICOTRANS ® Red L 2715 D) Base Stock Mixture 17.49

TABLE 7 Baseline Formulation Chamber Results Exhaust Results Temperature 2696 2097 (K) Pressure 100 12.1 (psia) Specific Heat (molar) Gas 11.31 11.23 Total 11.31 11.32 Molecular Weight 38.7 39.4 (grams/mole)

TABLE 8 Black Powder Chamber Results Exhaust Results Temperature 2132 1826 (K) Pressure 100 12.1 (psia) Specific Heat (molar) Gas 10.94 10.69 Total 17.58 17.71 Molecular Weight 59.2 59.8 (grams/mole)

Table 7 lists the expected performance of the propellant formulation at a chamber pressure of about 100 psia and Table 8 lists the corresponding expected performance of black powder as derived utilizing the PROPEP simulation software. The expected ballistic performance of black powder at various chamber pressures is presented in FIG. 5. FIGS. 4 and 5, show the respective Density Impulse (I_(d)), Specific Impulse (I_(sp) in lb/sec), Combustion Temperature (T in ° F.), and Gas Molecular Weight (MW) relative to chamber pressure (psia).

The burn rate of the propellant composition of an embodiment of the invention was characterized to be r=0.108P _(c) ^(0.3613).

At a chamber operating pressure of about 100 psia, the predicted molecular weight of the combustion mixture of the propellant formulation is about 38.7 grams per mole, which compares favorably relative to the predicted molecular weight of the combustion product of black powder of about 59.2 grams per mole.

Thus, compared to black powder, the propellant formulation of the present invention can provide a reduction in exhaust gas molecular weight of at least 20.3%. The resultant specific impulse was measured to be about 138.7 lbs/lbs-m/sec, which is about 47% greater than the specific impulse of black powder.

Example 9 Variation of Oxidizer to Fuel Ratio

Several modifications of the propellant formulation as substantially disclosed in Example 8 were evaluated. In particular, the effect on the specific impulse of varying the ratio of potassium perchlorate to dicyandiamide was evaluated utilizing the widely available propellant evaluation program, PROPEP code.

The predicted specific impulse at a chamber pressure of about 100 psia is presented in FIG. 6, which shows the propellant formulation of the invention can provide a specific impulse of at least 145 lb/sec for potassium perchlorate/dicyandiamide weight ratios from about 64:23.2 to about 71:15 and that a peak specific impulse of about 148 lb/sec can be achieved at a potassium perchlorate/dicyandiamide weight ratio of about 68/19.2.

Notably, the peak ratio is not stoichiometrically balanced; but can be considered to be underoxidized. This condition, it is believed, leads to reducing, if not minimizing, the average molecular weight of the exhaust gases. That is, the reducing atmosphere in the chamber can provide species that are not converted to the oxidized state. These conditions facilitate utilization of paper casing when exposed to the combustion gases at a temperature of about 4438° F. Specifically, the exhaust of the propellant formulation of the invention is expected to comprise about 0.000386 moles of free oxygen at a chamber pressure of about 100 psia, which is considerably less than the corresponding oxygen concentration of about 0.349 moles of black powder exhaust. Indeed, if free oxygen were present, oxidizable casings, such as paper, would degrade, e.g. burn, to failure. In the absence of oxygen, however, paper casings would pyrolyze to form a layer of insulating char that protects unexposed casing material.

Example 10 Effect of Binder Content

Further modifications of the propellant formulation as substantially disclosed in Example 8 were evaluated. In particular, the amount of vinyl resin (VROH) and modifier (HELOXY™ 505 modifier) was evaluated.

Increasing the amount of binder would improve pressability of the propellant composition into propellant grains as well as reduce the sensitivity thereof to friction, impact, electrostatic discharge, blasting cap; a reduction in specific impulse would likely result, as illustrated in FIG. 7. In particular, FIG. 7 illustrates predicted specific impulse (in lb/sec) relative to the chamber pressure with about 5 wt % binder content (indicated as reference 5), with about 7 wt % binder content (indicated as reference 7), and with about 10 wt % binder content (indicated as reference 10).

Example 11 Effect of Catalyst Content

Further modifications directed to the catalyst content of the propellant formulation as substantially disclosed in Example 8 were evaluated.

Although, it is believed, that specific impulse and density is typically determined by the combination of the oxidizer (e.g., potassium perchlorate), fuel (e.g., dicyandiamide), and the binder (e.g., vinyl resin/epoxidized castor oil), the burn rate catalyst may dictate other characteristics and stability of the sustained combustion at low, less than about 500 psia, or even atmospheric pressures. In particular, the selected burn rate catalyst of the invention can reduce or modify the high burn rate exponent (typically about 0.7) associated with potassium perchlorate while improving the burn rate coefficient.

For example, the effect of the amount of SICOTRANS® Red L 2715 D is presented in FIG. 8. In particular, the burn rate at about 2 wt % (indicated as reference 2), at about 4 wt % (indicated as reference 4), at about 6 wt % (indicated as reference 6), and at about 8 wt % (indicated as reference 8) of SICOTRANS® Red L 2715 D iron oxide catalyst relative to chamber pressure is shown.

Example 12 Effect of Oxidizer and/or Fuel Component

Further modifications directed to the catalyst content of the propellant formulation as substantially disclosed in Example 8 were evaluated.

In particular, the propellant formulation as substantially disclosed in Example 8 was modified to replace a portion of the potassium perchlorate with guanidine nitrate resulting in about 50.04 wt % potassium perchlorate and about 22 wt % guanidine nitrate. The specific impulse at a chamber pressure of about 100 psia is predicted to be about 145 lb/sec. The ballistic performance relative to chamber pressure is presented in FIG. 9.

The propellant formulation as substantially disclosed in Example 8 was also modified to replace dicyandiamide with triaminoguanidine (about 31.58 wt %). The specific impulse at a chamber pressure of about 100 psia is predicted to be about 166 lb/sec, which may be advantageous in highly energetic applications such as, but not limited to, inflators, guillotine cutters, power cartridges, bursting charges, expelling charges, pin pullers or pushers, bellows motors, signal rockets, and explosive actuated switches or valves. The ballistic performance relative to chamber pressure is presented in FIG. 10.

The propellant formulation as substantially disclosed in Example 8 was also modified to replace dicyandiamide with a comparable amount of hexamethylene tetramine (about 14.22 wt %), which may be advantageous in highly energetic applications. The balance of the formulation comprised of about 73.04 wt % potassium perchlorate, about 5.66 wt % iron oxide catalyst as SICOTRANS® Red L 2715 D, about 4.16 wt % VROH resin, and about 2.92 wt % functionalized modifier as HELOXY™ 505. The specific impulse at a chamber pressure of about 100 psia is predicted to be about 153 lb/sec. The ballistic performance relative to chamber pressure is presented in FIG. 11.

INCORPORATION BY REFERENCE

The entire contents of all patents, published patent applications, and other references cited herein are hereby expressly incorporated herein in their entireties by reference. Pending U.S. patent application Ser. No. 10/295,308, entitled COMPOSITE PROPELLANT COMPOSITIONS, filed Nov. 14, 2002, is incorporated herein by reference in its entirety.

Having now described some illustrative embodiments of the invention, it should be apparent to those skilled in the art that the foregoing is merely illustrative and not limiting, having been presented by way of example only. Numerous modifications and other embodiments are within the scope of one of ordinary skill in the art and are contemplated as falling within the scope of the invention. In particular, although many of the examples presented herein involve specific combinations of method acts or system elements, it should be understood that those acts and those elements may be combined in other ways to accomplish the same objectives.

Further, acts, elements, and features discussed only in connection with one embodiment are not intended to be excluded from a similar role in other embodiments.

It is to be appreciated that various alterations, modifications, and improvements can readily occur to those skilled in the art and that such alterations, modifications, and improvements are intended to be part of the disclosure and within the spirit and scope of the invention.

Moreover, it should also be appreciated that the invention is directed to each feature, system, subsystem, or technique described herein and any combination of two or more features, systems, subsystems, or techniques described herein and any combination of two or more features, systems, subsystems, and/or methods, if such features, systems, subsystems, and techniques are not mutually inconsistent, is considered to be within the scope of the invention as embodied in the claims.

Use of ordinal terms such as “first,” “second,” “third,” and the like in the claims to modify a claim element does not by itself connote any priority, precedence, or order of one claim element over another or the temporal order in which acts of a method are performed, but are used merely as labels to distinguish one claim element having a certain name from another element having a same name (but for use of the ordinal term) to distinguish the claim elements.

Those skilled in the art should appreciate that the parameters and configurations described herein are exemplary and that actual parameters and/or configurations will depend on the specific application in which the systems and techniques of the invention are used. Those skilled in the art should also recognize or be able to ascertain, using no more than routine experimentation, equivalents to the specific embodiments of the invention. It is therefore to be understood that the embodiments described herein are presented by way of example only and that, within the scope of the appended claims and equivalents thereto; the invention may be practiced otherwise than as specifically described. 

1. A rocket engine comprising: a case having a tensile strength of less than about 5,000 KPa; and a composite propellant grain disposed in the case, the composite propellant grain comprising an oxidizer, a burn rate catalyst, and a binder comprising a vinyl resin, and a polyfunctional ether.
 2. The rocket engine of claim 1, wherein the burn rate catalyst is acicularly-shaped.
 3. The rocket engine of claim 2, wherein the burn rate catalyst comprises iron oxide.
 4. The rocket engine of claim 1, wherein the oxidizer comprises potassium perchlorate.
 5. The rocket engine of claim 1, wherein the polyfunctional ether comprises a triglycidyl ether.
 6. The rocket engine of claim 5, wherein the triglycidyl ether comprises a polyepoxide resin.
 7. The rocket engine of claim 6, wherein the binder further comprises an amide.
 8. The rocket engine of claim 7, wherein the amide comprises dicyandiamide.
 9. The rocket engine of claim 1, wherein the vinyl resin comprises a hydroxyl-functional polymer selected from the group consisting of vinylchloride, vinylacetate, hydroxyalkylacetate, and mixtures thereof.
 10. The rocket engine of claim 1, wherein the case has a bursting strength of less than about 1,000 KPa.
 11. The rocket engine of claim 1, wherein the case comprises paper or paperboard.
 12. The rocket engine of claim 1, wherein the case is comprised of a material that is free of metals or metal alloys.
 13. The rocket engine of claim 1, wherein the composite propellant grain comprises: an oxidizer consisting essentially of potassium perchlorate; a burn rate catalyst consisting essentially of acicularly-shaped iron oxide nanoparticles; and a binder consisting essentially of a vinyl resin and a triglycidyl ether cross-linked with dicyandiamide.
 14. A method of preparing a pressable composite propellant formulation comprising: mixing a vinyl resin, a vegetable oil, and a solvent to provide a binder mixture; adding a burn rate catalyst to the binder mixture; adding an amide to the binder mixture; and reducing the solvent from the binder mixture to about 7 wt % to provide a mixed formulation.
 15. The method of claim 14, wherein the vegetable oil comprises epoxidized castor oil.
 16. The method of claim 14, wherein the vinyl resin comprises a hydroxyl-functional polymer selected from the group consisting of vinylchloride, vinylacetate, hydroxyalkylacetate, and mixtures thereof.
 17. The method of claim 14, wherein the burn rate catalyst comprises acicularly-shaped iron oxide nanoparticles.
 18. The method of claim 14, wherein the amide comprises dicyandiamide.
 19. The method of claim 14, wherein the solvent comprises an ester.
 20. The method of claim 19, wherein the solvent comprises ethyl acetate.
 21. The method of claim 14, further comprising an act of drying the burn rate catalyst prior to performing the act of adding the burn rate catalyst to the binder mixture.
 22. The method of claim 14, further comprising an act of extruding the mixed formulation to form pellets thereof.
 23. The method of claim 22, further comprising an act of drying the pellets to reduce the amount of solvent present therein.
 24. The method of claim 23, further comprising an act of encasing at least one pellet in a case comprising a case material having a tensile strength of less than about 2,000 KPa.
 25. The method of claim 24, wherein the case material has a specific gravity of less than about
 2. 